Compound ram jet turbo-rocket engines



Y 2, 1961 B. L. BELCHER ErAL 2,982,093 oMPouND RAM JET TURBO-ROCKETENGINES 7 Sheets-Sheet 1 Filed Aug. 26, 1958 /NVENTORS May 2, 1961 B.L.. BELCHER ETAL 2,982,093

COMPOUND RAM JET TURBO-ROCKET ENGINES Filed Aug. 26, 1958 '7Sheets-Sheet 2 BY www) ATTORNEY:

Filed Aug. 26,l 1958 May 26.1961 B. L. BELCHER Erm. 2,982,093

COMPOUND RAM JET TURBO-ROCKET ENGINES 7 Sheets-Sheet 3 76 \73 77) l aoArron/vf May 2, 1961 B. L.. BELCHER ET AL 2,982,093

COMPOUND RAM JET TURBO-ROCKET ENGINES 7 Sheets-Sheet 4 Filed Aug. 26,1958 [infill/1111]] Q, v V m n,

lllllll llnllllil-lllsl May 2, 1961 B. L. Bl-:LcHER ETAL 2,982,093

COMPOUND RAM JET TURBO-ROCKET ENGINES Filed Aug. 26, 1958 7 Sheets-Sheet5 @mil @ywl uw ATTORNEY May 2, 1961 B. L.. BELCHER ErAL COMPOUND RAM JETTURBO-ROCKET ENGINES Filed Aug. 26, 1958 7 Sheets-Sheet 6 V BLYTHATTORNEY RONALD MOP/27S GEORGE F. UMD/V.

May 2,1961 B. l.. BELCHER ET AL 2,982,093

coMPoUND RAM JET TURBO-ROCKET ENGINES Filed Aug. 26, 1958 7 Sheets-Sheet'7 NQ .0.2 PQ

2,982,093 Patented May 2, 1961 COMPOUND RAM JET TURBO-RUCKET ENGINESBryan Leslie Belcher, Jack Vallis Blyth, David Edwin James Buckingham,Alan Leslie Davies, Reginald Henry Douglas Chamberlin, Alan Leslie RoyFletcher,

Ronald Edward Morris, Geoffrey Charley Gerald Mansiield, and George'Frank Upton, all of London, Eng land, assignors to D. Napier & SonLimited, L'ondon, England, a company of Great Britain Filed Aug. 26,1958, Ser. No. 757,404

Claims priority, application Great Britain Aug. 3i), 1957 5 Claims. (Cl.60-35.6)

This invention relates to compound ram jet turborocket engines of thetype including an air intake at the forward end of the engine, a maincombustion chamber to which fuel is supplied and an exhaust pass-ageterminating in a propulsion nozzle, and a rocket type gas generatorarranged to drive a turbine which is in turn coupled to a compressor,the compressor^ being arranged to receive air from the air intake and todeliver it to the main combustion chamber.

Such engines operate under two different regimes. In the ram jet regimeair is admitted to the engine through the air intake at the forward endand is compressed by the ram effect and supplied under pressure to themain combustion chamber where the fuel is burnt. In the rocket regimethe rocket `drives the turbine which in turn drives the compressor andthe air supplied to the main combustion chamber is pressurised by thecompressor. Thus the rocket regime is particularly adaptable tooperations at forward flight velocities which are insufficient toprovide the necessary ram eect. In some operating conditions it may bedesirable to operate the engine on both systems simultaneously. In anycase it will be seen that the internal operating conditions of theengine are subject to wide fluctuations due to the wide range in forwardflight velocities `and also to Vthe various different regimes in whichthe engine operates. In particular the operating conditions, such aspressure, velocity and temperature of the 4air passingthrough the airintake to the main combustion chamber, are liableto fluctuateconsiderably. It is an object of the present invention to provide anengine of the kind referred to which Will be capable of eicientoperation over the different conditions that may be expected.

According to theV present invention, therefore, a compound ram jetturbo-rocket engine of the kind referred to includes a by-pass passagebetween the air intake and the main combustion chamber lb'y-passing theworking passages of the compressor.

The engine also preferably includes valve mechanism arranged to controlthe flow of air through the by-pas passage;

' According to a preferred feature of the invention the compressorincludes lat least two rings of rotor blades and at least two rings ofstator blades downstreamvthereof, the downstream ring of stator bladesbeing adjustable, each blade on a pivotal axis which is substantiallyradial to the axis of rotation of the compressor, and includingadjusting mechanism arranged to pivot each blade of this downstream ringinto one or other of two operatingpositions, in one of which the bladering operates at maximum .efficiency as a normal stator blade ring,while in the other position the blades are positioned toV give themaximum effective Vthroat area between blades.

In such case according to another preferred feature of the invention thevalve mechanism controlling the air flow through the by-pass passage ispreferablyassociated with and arrangedrto act simultaneously withthentcd States Patent mechanism adjusting the position of the stator bladesof the compressor.

According to yet another preferred -feature of the invention the Iairintake comprises two series of circumferentially spaced flaps arrangedto open or close apertures in the outer wall thereof, the firstupstreamseries of aaps being pivotally mounted at their rear ends whilethe second downstream series are pivotally mounted at their forwardends, both series of flaps being arranged to open outwards.

It will be` understood that to obtain the maximum rate of air flowthrough the air intake to the combustion chamber it is desirable thatair should pass through the compressor in addition to passing throughthe by-pass passage. Thus according to another preferred Ifeature of theinvention the compressor rotor and the turbine rotor are coupled to oneanother through a unidirectional clutch arranged to enable thecompressor to over-run the turbine or to freewheel or Windmill when theturbine is stationary.

The exhaust nozzlev assembly of the engine must also be capable ofecient operation over the full range of flight conditions, and accordingto yet another preferred feature of the invention the exhaust nozzleassembly comprises an 4outer venturi shaped Wall and an innerdouble-tapered central bullet (that is to say tapered in both directionsfrom the centre), the outer Wall or the central bullet `being movable ina longitudinal direction relative to one another between two mainpositions in one of which the maximum diameter of the bullet is adjacentthe point of minimum diameter of the venturi throat, While in the othermain position the maximum diameter of the bullet is adjacent the rearend of the outer wall.

The fuel supply system for the engine preferably comprises a rocket fuelpump driven by the rocket turbine, and supplying fuel to the rocket gasgenerator, and a separate main fuel pump capable of being driven`alternatively by the vrocket turbineor by an independentmotor andsupplying fuel to the main combustion chamber.

In such an engine including component or ,auxiliary mechanisms such aslubrication or hydraulic servo systems which necessarily contain aIliquid medium such as oil, the fluid' circuits of the mechanisms arepreferably yarranged in at least two separate closed circuits eachinconporating a separate heat exchanger,A the main liquid lfuel supplyto the Vengine beingpassed in series through these heat exchangers tomaintain the liquid in one of the circuits at the minimum temperatureattainable by the cooling effect of the fuel at its delivery temperatureto the engine.

The invention may be performed in various different ways, but onespecific embodiment will now be described by way of example Vas appliedto an aircraft propulsion engine which is illustrated in sectionalelevation in Figures 1A and 1B of theaccompanying drawings, of which:

Figure 2 is a diagrammatic illustration of the fuel lubrication andhydraulic servo systems; f

Figure 3 s a diagrammatic illustration of the driving mechanism for thepumps;

Figure 4 is a sectional side elevation on an enlarged'W scale of thecompressor;

' Figure 5 is a diagrammatic developed sectional view of the compressorblading;

Figure 6 is Ia sectional elevation on an enlarged scale through thefreewheel mechanism 28 between the turbine yand compressor;

Figure 7 is Ia diagrammatic sectional elevation on a slightlyY reducedscale of the air intake of therengine showing ,the pressure sensingheads; i

iFigure 8 is a diagrammatic illustration of the automatic servo controlmechanism controlling the airintake lla-ps;

Figure 9 is a diagrammatic illustration of the automatic controlmechanism for the spill flaps; and p Figure lO is a graph illustratingvarious conditions which occur in the air flow enteringthe air intake.

This engine is a compounded ram jet-turbo rocket engine. It includes anouter generally cylindrical casing 10, the front end 11 of which formsvwith a conical centre body 12 an annular air intake 13 leading to anaxial diffuser passage 14 and then to an axial flow compressor 15 fromwhich the air passes rearwardly through an annular air duct 16 into anannular combustion chamber 17 to which a hydrocarbon fuel suchaskerosene is supplied through burners 18. The hot productsL ofcombustion issue through a nozzle 19 at the rear ofthe engine as a highspeed propulsion jet. A bullet 20 is provided for adjusting the area ofthe nozzle, this bullet being axially movable by a hydraulic ram 211disposed in the-front part of the engine and connected to theV bullet bya long shaft 22 extending rearwardly through the centre of the engine.The ram 21 is controlled bya follow-up servo valve 23.

The compressor comprises two rotor blade rings 25, 26 mounted on ahollow shaft 27 which is connected by freewheel mechanism 28 to anotherhollow shaft 29 at the downstream end of which is mounted a two stageaxial flow turbine 30. This turbine is driven by a rocket systemincluding vcatalytic decomposition chambers 31 to which hydrogenperoxide is supplied and decomposed to form oxygen-rich steam, androcket combustion chambers 32 into which some paraffin is thenintroduced for combustion with some of the oxygen. The combustionproducts, which are still rich in oxygen, are expanded in the turbine30, and then pass through ducts 33 to enter the main combustion chamber17 along with the air from the said annular duct 16, to contribute tothe propulsive effect of the jet. Also, the residual oxygen content ofthe turbine eiuent assists the main combustion process.

Provision is also made for the engine to operate as a pure ram jet whena sufhciently high speed has been reached. For this purpose anv annularby-pass duct 35 is provided around the axial flow compressor 15 and thedownstream end of this by-pass duct communicates through flap valves 36,pivoted at `37, with the said annular air duct 16 at the downstream endof the compressor. These flap valves 36 are actuated by hydraulic servomotors 38 which also actuate a ring of adjustable stator blades 39 forthe last stage of the compressor, to increase the flow through thecompressor andl reduce losses therein during ram jet operation. At apredetermined Mach number, when the ram effect alonewill providesufficient compression of the air, and which is sensed by a Mach meterindicated generally at 34, the hydrogen peroxide and paraffin supply tothe rocket system isautomatically shut off, thus stopping the turbine 30and removing the power supply for the axial flow compressor `15. At thesame time the by-pass flap valves 36 are opened to allow air from theair intake 13 to pass through the annular by-pass duct 35 to thecombustion chamber 17. The free wheel mechanism 28 between the turbine30 and the axial ow compressor 15 permitsrthe latter to windmill duringram jet operation.

The centre body `12 has a conical nose tip 40, and cooperates with theouter cowl 11 to provide shock compression during ram jet operation.Behind the lip of the cowl at the beginning of the diffuser section 14there is an annular port providing communication between the outside ofthe engine and the diffuser section. During ram jet operation this portis closed by a series of petal type aps 41 pivotally attached to theouter shell of the engine at their rear ends 42. During turbo rocketoperation these flaps are swung outwards as shown in chain lines by ahydraulic servo motor 4'3 controlled by a follow-up servo valve 46, soas to provide an additional annular air intake of larger diametersurrounding the main air intake 13.

Towards the rear end of the diffuser section 14 and upstream of theaxial ow compressor 15 there is arranged another series of similarpivoted flaps 44 controlling another annular port inthe outer shell I11which constitutes a controllable spill-*port which is-opened during ramjet operation to the extent required to maintain the desired shockpattern at the intake. These flags 44 are pivotally connected to theshell at their upstream ends 45 and are opened outwards by a hydraulicservo motor 47 controlled by a follow-up servo valve 48.

Thermally insulated compartments are provided in the engine within theannular air duct through the engine. There is a front compartment 50lying mainly within the diffuser section `14 and terminating adjacentthe forward end of the vcompressor 15. This compartment houses thehydraulic servomotors 43, 47 which are connected through mechanicallinkages with the said two series of flaps 41 and 44, and their controlvalves 46, 48. It also houses the Mach meter 34 and the shock sensingsystem, a forward bearing 51 for the axial flow compressor 15, and thehydraulic ram 21 for adjusting the position of the nozzle bullet 20.

A second insulated compartment 52 surrounds the shaft 29 connecting theturbine to the compressor and encloses the freewheel mechanism 28, pumpsfor lubricant, servo motor fluid, parain and hydrogen peroxide, gearingfor driving these pumps and other auxiliaries. This compartment alsocontains a sump of lubricating oil, two heat exchangers through whichparaffin fuel is pumped in succession as a coolant and a metering unitfor the paraffin.

The fuel supply system and lubricating and hydraulic servo supplycircuits are illustrated in Figure 2.

Paran is supplied from a tank 55 mounted outside the engine, for examplein an aircraft or missile to which the engine is attached. The fuelin-this tank is used as a coolant or heat sink for parts of the aircraftand a heat exchanger 56 and fuel circulating pump 57 are provided forthe purpose. The fuel in the tank 55 may thus be at a relatively hightemperature possibly close to the vaporisation point. In order-to permitthe fue] to be further heated, and used as a coolant in the process, itis admitted to the engine through a pressurising pump 58 which raisesits pressure substantially. The fuel then passes in succession throughtwo heat exchangers 59, 6i).

In the first heat exchanger 59 the paraffin takes up heat fromlubricating oil which is delivered by a pump 63 from a sump 62, throughthe heat exchanger, to the gearing for the pumps and auxiliariesindicated diagrammatically at 64 and to the turbine and compressorbearings indicated at 65 all these parts being within the said insulatedcompartments 52 and therefore receiving little heat. In the second heatexchanger the parain takes up heat from the actuating fluid of theseveral hydraulic servomotors such as 43, 47, 21 and'38. The parafn thenpasses to a metering unit 67 and on to the burners 18 via ducts 68.

The hydraulic servo circuit includes a pressurising pump 66 whichimpells the servo uid through the heat exchanger 60, and thence inparallel to the servo rams two of which are indicated. diagrammaticallyat 38 and 43.

The heat exchanger 59'being upstream of heat exchanger 60 in the fuelsupply line, will have a lower temperature datum, and this is desirablesince the lubricating oil may tend to carbonise if heated excessivelywith risk of damage to bearings and other high speed parts. Thehydraulic servo liquid' on Ithe other hand can safely be allowed toreach higher temperatures, since the servo mechanisms are relativelyslow moving. In the present example the paraffin will enter the engineat pump 58 at a temperature not exceedingk 150 C. and in passing throughthe two heat exchangers its temperature may be raised 30 C. It thenpasses to afuel metering unit 67, and thence to delivery liners 68leading to burners 18 in the main4 combustion chamber-17. Thelubricating oil delivers it under pressureto the catalytic decompositionchambers 31 via a conduit 75. The turbine shaft 29 is also arranged todrive through the gearing '71, 72 and further gearing 76, three units77, 66 and 69; Unit'77 s a hydraulic r.p.m. signal generator whichprovides a fluid pressure proportional to the speed of rotation of theturbine, and is used for automatic control purposes, Servo p ump 66 isarranged to provide servo fluid under pressure to operate the servos 43,47, 38 etc. when the rocket turbine 30 is in operation. Pump 69 isarranged to receive parain fuel from the main aircraft fuel tank 55through duct 78 via a filter 79 and to deliver the fuel underincreasedpressure to the rocket combustion chambers 32 via a conduit 80.4

The gearing 7,6 is connected ,through a free wheel device 82 to anothergear train indicated at 83, this gearing being arranged to drive thesecond servo pump 84, as illustrated in Figure 2, and the ramjetparaffin pump 58, which delivers parain at high pressure through duct-87to the heat exchangers 59, 60, and thence to themain burners 18. Thisgear .train 83 also drives pumps 63, und 61, the pressure and scavengepumps in the circuit illustrated in Figure 2.

The gearing 83 is also connected by a shaft 90 toa radial flow turbine91 to which air is supplied under pressure from the annular air d uct16. The exhaust of this turbine is allowed to escape to a low pressureregion such as atmosphere through a duct 92 in which is arranged anadjustable throttle valve 93 under the control of a servo control device94.

Thus in operation when the engine is first started by an externalbooster pump or starter motor, the rocket turbine 30 will be driven bythe rocket exhaust and this will drive both the -main compressor throughthe fr'ee wheel 28, and also all units driven by gearing 76, and inaddition will drive through the free Wheel 82 and gearing 83 all theunits connected to this gearing 83. When the engine has achieved aforward flight speed suiiicient for ram jet operation the rocket sectionwill be shut down 4and the turbine 30 will stop. The main compressor -15can then windmill overrunning the turbine shaft by means of the freewheel :28, but the gearing 76 will not be driven by the turbine and theunits 73, 77, 66 and 69 will be out of operation. The air turbine 91however will then be driven by the relatively high pressure in theannular air duct 16 and through the shaft 90 and gearing 83 all unitsassociated therewith will be driven. In particular pump 58 supplyingfuel to the main burners 18, servo pump 84 and lubrication pumps 61 and63 will then continue to be driven by the air turbine. Y

'Ihe compressor 15itself is of the two stage type including the tworotor blade rings 25, 26 mounted on rotor discs '101, 102, attached tothe compressorshaft 27. Between the rotor blade rings there is mounted aring of stator, blades 103 which are angularly fixed and non-adjustable.Downsteam of the second rotor blade ring 26 there isf mounted a secondring of stator blades 39 which are each angularly adjustable about axesthrough each blade which are radial to the main axis of the compressor.A further series of iixed straightener vanes 105 is provided downstreamof the second adjustable stator blade ring 39. p

The blades 39 are each connected rigidlyto a pin 106 'capable ofrotating in a bushing carried by a fixed part of the engine and theinner end of this pin is connected to an'otset crank 107 which is inturn connected through pivoted links 108 and 109 to a point on a bellcrank lever 110 which is angularly adjustable by means of the doubleacting servo ram 38. The ram comprises a ram piston l111 connected to apiston rod passing through a gland at one end of the ram cylinder andmeans are also provided for admitting pressure fluid to either end ofthe ram cylinder as required under the `control of an automatic servovalve. The servo valve itself forms no part of the present invention'and will not therefore be described in detail.` The hydraulic servo ram38 is also arranged toactuate the flap valves 36. To this end the bellcrank lever 110 is connected through a pivoted link 112 to a point latth'e upstream end of each flap valve. In the position illustrated in theupstream end of the ap valve abuts against a cylindrical shield 113which separates the by-pass passage 35 from the outlet passage of thecompressor, both these passages communicating at their downstream endswith the annular air passage 16. In this position the liap valve 36closes the by-pass passage 35 and air can only pass into the passage 16via the compressor'. In the other operative position of the flap valve36 as illustrated in chain lines the by-pass passage 35 is opened to thepassage 16 and air can then ow both through the compressor and throughthe =bypass passage.

In operation when the rocket turbine 30 is driving the compressor 15 theservo ram piston 111 is in the position illustrated and the flap valve36 is closed preventing air ilow through the by-pass passage, and thesecond row of stator blades 39 are in the position` indicated in fulllines in Figure 5. The compressor thus delivers air to the maincombustion chamber 17 and operates normally at full eiciency.

When theengine has attained a forward iiight velocity at whichthe rampressure is suiiicient to support combustion the supply of -fuel to therocket engine driving Vthe turbine is shut off automatically or by thepilot selection and the turbine stops. At the same time the servo valveautomatically reverses the high pressure connections to the ram cylinder38 and the ram piston is moved to the right in Figure 2 thus opening theiiap valves 36 so that the by-pass passage 35 allows air to ow aroundthe compressor into the air passage 16, and at the same timetheadjustable stator blades 39 are rotated into their second operativepositions as indicated in chain lines in Figure 5. In thisposition ofthe blades the maximum throat area is achieved with the minimumresistance to iiow and the turbine rotor blade rings 25, 26 willwindmill `under the air flowing through the compressor thus augmentingthe total air flow into the main combustion chambers 17.

The servo valve controlling actuation of the ram 38 may be arranged tobe responsive to the speed of rotation of the turbine. Thus at allturbine speeds above a predetermined gure the servo valve will bearranged to hold the ram piston 111 in the position illustrated, whilewhen the turbine speed falls below this valve as a result of the fuelsupply being shut off to this rocket combustion chambers the ram pistonwill move into its alternative operative position.-

The freewheel mechanism 28 is illustrated in detail in Figure 6. Theturbine shaft 29 is formed with a hollow cylindrical extension 1351 atits end adjacent the compressor shaft 27 and the external surface ofthis extension piece is provided with a quick pitch helical screwthread132. Cooperating with this screwthread is an internal screwthread sleeve133 which constitutes a movable-clutch member and which is formed with aring of dogs 134 which can be engaged with a corresponding -ring -ofdogs 135 formed on a ange secured to the compressor shaft 27. The sleeve133 is also provided with an inwardly projecting radial ange 136 whichin the position shown abuts against a locking ring 137 while in theother limiting position of the sleeve this ange abuts against a shoulder138 formed on the turbine shaft. The sleeve 133 is also formed with anoutwardly projecting radial flange 139to which are connecteda series ofresilient friction leads140`which'bear against the face of a flange 141formed ofi aV sleeve 142 which is splined or keyed to the compressorshaft 27. This sleeve 142 is urged towards the turbine shaft by acompression spring 143.

In the position illustrated in Figure 6 the dogs 134, 135 are out ofengagement and the mechanism-is in its free wheeling position. Thecompressor shaft 27 can rotate in a clockwise direction when viewed fromthe turbine and there will be a slight frictional drag between thefriction leads 140 and the ange 141.

If the turbine shaft 29 is driven in this same direction of rotation ata speed faster than the compressor shaft 27 the frictional drag on theleads will cause the sleeve 133 to rotate on the quick pitch screwthread132 and the sleeve will move axially towards the compressor until thedogsl 134, 135 engage and ultimately the flange 136 on the sleeve abutsagainst the should 138. During this axial movement of the sleeve aspring 143 will be compressed thus increasing the loading pressurebetween the leads 140 and the tiange 141 and so increasing thefrictional drag and hence the torque available for rotating the sleeve133. When the dogs 134, 135 are fully engaged the turbine shaft 29 willthen drive the compressor shaft 27.

The two sets of flaps 41 and 44 can be arranged to be controlledautomatically by the apparatus illustrated in Figures 7, 8 and 9. Theapparatus comprises a series of pressure sensing heads as illustrateddiagrammatically in Figure 7. A first pressure head 150 is mounted onthe inner surface of the cowl 11 at a point adjacent to but a shortdistance downstream of the forward lip 151 of this cowl. A secondpressure sensing head 152 is positioned on the surface of the conicalcentre body 40 at a point some distance downstream of the extreme tip orapex of the cone. It will be seen that the cone is in fact of doublefrusto-conical form and includes a first sharp angled conical section153 and a rearward frusto-conical section 154, the pressure head 152being situated somewhat upstream of the junction between these twosections. At the extreme forward end or point of the conical centre bodythere is provided a pivot head on a forl' wardly facing probe comprisingan inner tube 155 with a sharp intake lip 156 at its forward end, and asurrounding hollow tube 157 provided with side apertures 158. Thepressure sensing heads 150, 152, 156 and 158 are connected respectivelyto pressure conduits `159, 160, 161 and 162.

At low flight speeds below Mach .95 the spill flaps 44 remain closed.Within the range of flight speeds the intake flaps 41 are under thecontrol of the valve apparatus 46 which is illustrated in detail inFigure 8. This valve r' is influenced by the pressures in the conduits161 and 162, that is to say by the total pressure at the pitot head 156and the ambient air pressure as sensed by the side orices 158. Theapparatus comprises a piston type control valve 165 arranged in a valvecylinder 166 and provided with lands 167, 168 which control the ow of ahigh pressure servo tiuid from an inlet aperture 169 selectively to oneor other of the delivery passages 170 and 171. These delivery passagesare connected to opposite ends of the ram cylinder 172 which houses thereciprocating ram 43, this ram being connected by a mechanical linkageto the intake flaps 4l (one flap only being illustrated in Figure 8).The opposite ends of the valve chamber 173 are closed by exiblediaphragms 174 and 175 and these two ends of the chamber are connectedto a low pressure fluid line 176. It will be seen that when one end ofthe ram cylinder 172 is connected to the high pressure fluid line 169the opposite end of the ram cylinder is connected to the low pressureline 176 and vice versa.

The valve apparatus also includes chambers lying on the. remote sides ofthe diaphragms 174 and 175. One of these. chambers is itself sub-dividedby a flexible diaphragm 177 the centre part of which is rigidlyconnected to the valve stem 165. The upper part of this sub-dividedchamber is connected to the pressure conduit 161 which leads to theforward facing orifice 156 of the pitot tube, while the lower part ofthis chamber is connected to the conduit 162 which leads to the sideorifices 158 on the pitot head. The chamber 178 at the opposite 4end ofthe valve is also connected to the conduit 162 and this chamber containsan evacuated flexible capsule 179, one end of which is rigidly connectedby a bolt 180 to the end of the valve chamber, while the other end ofthe diaphragm is rigidly connected to the valve stem 165.

`It will be seen that the servo fluid pressure acting on the valve stemis substantially balanced and the total force exerted on the valve stemis determined by the resilience of the bellows and the pressures in theconduits 161 and 162 and by the effective area of the capsule 179 and ofthe diaphragms and 177. In practice these areas and the spring rate ofthe diaphragm are so selected that the valve stern Will move to causethe ram 43 to be urged downwards to hold the intake flap 41 in its openposition, until the total pressure in the conduit 161 exceeds that inthe conduit 162 by a predetermined value. This value, which is derivedfrom the pitot head, is in effect the Rayleigh number of the air flowconditions due to the forward Iflight velocity and the dimensions of thcvarious parts are so selected that the valve will operate in this mannerwhen the forward flight speed reaches a valueof Mach 0.95. At this ightspeed the further automatic Mach No. control 34 comes into operation toout off the supply of fuel to the rocket engine and at all flight speedsabove this figure the engine then operates as a ram jet. At all suchhigher flight speeds the intake flaps 41 remain fully closed under theinfluence of the valve 46.

At higher supersonic ight speeds a shock wave pattern exists in theneighbourhood of the air-intake which varies at different speeds butwhich adopts a pattern as illustrated in chain lines in Figure 7 whenthe engine is operating at the design point at approximately critical"conditions. This pattern comprises a lirst oblique shock wave extendingrearwardly from the forward end of the cone 40 to a point adjacent thelip 151 ofthe cowling. A second oblique shock wave 181 originates at thejunction between the forward cone 153 and the rear cone 154 and thisshock wave also extends to a point adjacent the lip 151 of the cowling.A third normal shock wave 182 occurs within the air intake itself at apoint somewhat downstream of the lip 151. Referring to Figure 10 it willbe seen that the critical condition corresponds to the maximum pressurerecovery or intake eiciency while the sub-critical condition occurs whenthe mass ow is below the critical value. The critical mass flowrepresents the maximum mass air flow that will pass through the intakeand at supercritical conditions the pressure recovery or intakeetiiciency is progressively reduced. When related to the shock wavepattern illustrated in Figure 7 the main noticeable change in the shockpattern as conditions changed from sub-critical to super-critical is themovement ofthe normal third shock wave 182 forwardly or rearwardly. Atsub-critical conditions this normal shock wave 182 tends to move forwardand vice versa.

It will be seen that the irst pressure head 150 is situated at orclosely adjacent to the position of the third shock wave 182 duringcritical operation. Any movement of this shock wave upstream ordownstream will thus provide a significant change in the pressure in theconduit 159 and this pressure is used to control the attitude of thespill aps 44 through the valve mechanism 48 which is illustrated indetail in 'Figure 9.

This valve mechanism comprises a piston type servo control valve 185controlling the supply of servo pressure uid from a high pressure source186 selectively to either end of a ram cylinder 187 containing the rampiston 47. The opposite end of the ram cylinder is similarly connectedto one or other olf-the low pressure lines 188, 189. The ram piston 47is connected by a mechanical linkage to the spillaps 44 (one only beingillustrated in its closed position in Figure 9).

The valve stem 185 extends through one end of the valve chamber and isconnected at pivotal joint 190 to one end of a oating beam 191, thispivotal joint 190 also being connected to the mid-point of a cxiblediaphragm 192 which sub-divides a chamber 193, the upper end of thischamber being connected to the conduit 159 while the lower part isconnected to the conduit 160. The beam 191 bears at an intermediatepoint in its length on a roller 194 which constitutes a movable fulcrumand is adjustable longitudinally by a rod 195 runder the control of thepilot or automatically by sensing of Mach No. The opposite end of thebeam 191 is connected to a vertical plunger 196 the lupper end of whichis carried in a sliding bea-ring in the valve chamber, while the lowerend is connected to an evacuated ilexible capsule or bellows 197. Thelower end of this bellows is connected by a bolt 198 to the lower end ofthe surrounding valve chamber. The interior of this valve chamber isconnected to conduit 162 and is thus subject to ambient static pressurewhich provides through the capsule a corresponding vertical force on theplunger 196. The diaphragm 192 is subjected on opposite sides, to thepressures in conduits 159 and 160 and is thus sensitive to changes inthe position of Ithe normal shock wave 182 and also to changes in thepressure behind the rst oblique shock wave 180. The resulting force actson the pivotal connection 190 at the opposite end of the beam 191 andthe differential resultant is transmitted to the valve piston 185, theactual value of this differential resultant being adjustable by means ofthe movable ulcrum 194.

In operation it will be seen that if the third shock wave 182 movesforwards as a result of the conditions becoming sub-critical thepressure in conduit 159 will be increased, which results in the valvepiston 185 moving downwardly and so causing the ram 47 to move to thetight imFigure 9 to move the spill aps 44 towards their open position asindicated in chain lines. Conversely a rearward movement ot the shockwave 182 will cause a decrease in the pressure in conduit 159 and thevalve 185 will move upwardly causing the ilaps 44 to move towards theirclosed position. In practice some instability of the shock wa-ve isbound to occur and the spill aps 44 will be continuously in movement,due to the pressure changes or -utters in the moving shock p-attern,tending at al1 times to maintain critical conditions in the air intake.

What we claim as our invention and desire to secure by Letters Patentis:

j 1. A compound ram jet turbo-rocket engine including an air intake atthe forward end of the engine, a main combustion chamber to which fuelis supplied and an exhaust passage communicating with said combustionchamber and terminating in a propulsion nozzle, Ia rocket type gasgenerator, a turbine arranged to be driven by the gas generated, and acompressor coupled to the turbine, the compressor being arranged toreceive air from the air intake and to deliver it to the main combustionchamber, and including a by-pass passage between the air intake and themain combustion chamber 4lay-passing the working passages of thecompressor and' a valve mechanism arranged to control the flow of airthrough the by-pass passage independently of the air ow through thecompressor.

2. A compound engine as claimed in claim 1, in which the compressorincludes at least two rings of rotor blades and at least two rings ofstator blades each downstream of one of the rotor blade rings, thedownstream ring of stator blades 'being adjustable, each blade on apivotal axis which is substantially radial to the axis of rotation ofthe compressor, and including adjusting mechanism arranged to pivot eachblade of this downstream ring into one or the other of two operatingpositions, in one of which the blade ring operates at maximum eiciencyas a normal stator blade ring, while in the other position the bladesare positioned to give the maximum effective throat area between blades.

3. A compound engine as claimed in claim 1, in which the air intakecomprises two series of circumferentially spaced aps a-rranged to openor close apertures in the outer wall thereof, the first upstream seriesof flaps being pivotally mounted at their rear ends while the seconddownstream series are pivotally mounted at their forward ends, bothseries of aps being arranged to open outwards.

4. A compound engine as claimed in clainrl, -in which the compressorrotor and the turbine rotor are coupled to one another through aunidirectional clutch arranged to enable the compressor to free-wheel orwindmill when the turbine is stationary.

5. A compound engine as claimed in cl-aim 2, including a couplinginterconnecting the said valve mechanism and the said blade adjustingmechanism to cause said adjusting mechanism to place the stator bladesin the position of maximum effective throat `area when the by-passpassage is opened by said valve mechanism.

References Cited in the tile of this patent UNITED STATES PATENTS2,582,848 Price Jan. 15, 1952 2,604,278 Johnson July 22, 1952 2,778,564Halford et al. Jan. 22, 1957 2,867,978 Peterson Jan. 13, 1959 2,873,576Lombard Feb. 17, 1959 2,896,408 ODonnell July 28, 1959 FOREIGN PATENTS749,009 Great Britain May 16, 1956

